System and method of fabricating and repairing a gas turbine component

ABSTRACT

A method of fabricating and repairing a gas turbine component having a plurality of cooling holes defined therein is provided. The method includes determining a parameter of a first cooling hole defined in the gas turbine component, and generating a tool path for forming a protective cap around the first cooling hole. The tool path is based at least partially on the parameter of the first cooling hole. The method also includes directing a robotic device to follow the tool path, and discharging successive layers of ceramic slurry towards the gas turbine component as the tool path is followed such that the protective cap is formed around the first cooling hole.

BACKGROUND

The present disclosure relates generally to gas turbine engines and,more specifically, to systems and methods of forming protective capsaround cooling holes in a gas turbine component.

In a gas turbine engine, air pressurized in a compressor is mixed withfuel in a combustor to generate hot combustion gases. Energy isinitially extracted from the gases in a high pressure turbine (HPT) thatpowers the compressor, and subsequently in a low pressure turbine (LPT)that powers a fan in a turbofan aircraft engine application, or powersan external shaft for marine and/or industrial applications. Generally,engine efficiency increases as the temperature of combustion gases isincreased. However, the increased gas temperature increases theoperating temperature of various components along the gas flowpath,which in turn increases the need for cooling such components tofacilitate extending their useful life.

For example, at least some known gas turbine components, such as blades,nozzles, and liners, require cooling during operation of the gas turbineengine. In at least some gas turbine engines, flowpath componentsexposed to hot combustion gases are cooled using compressor bleed air.For example, at least some known components channel the compressor bleedair through film cooling holes defined within the gas turbinecomponents. However, the gas turbine components generally have a limitedservice life and must be periodically serviced to ensure the gas turbinecomponents continue to function properly. Servicing the gas turbinecomponents typically includes removal of an existing thermal barriercoating and subsequent reapplication of a thermal barrier coating to thecomponents. The film cooling holes may become blocked when reapplyingthe thermal barrier coating, and cleaning and clearing the film coolingholes of the coating is a time-consuming and laborious task.

BRIEF DESCRIPTION

In one aspect, a method of fabricating or repairing a gas turbinecomponent having a plurality of cooling holes defined therein isprovided. The method includes determining a parameter of a first coolinghole defined in the gas turbine component, and generating a tool pathfor forming a protective cap around the first cooling hole. The toolpath is based at least partially on the parameter of the first coolinghole. The method also includes directing a robotic device to follow thetool path, and discharging successive layers of ceramic slurry towardsthe gas turbine component as the tool path is followed such that theprotective cap is formed around the first cooling hole.

In another aspect, a system for use in fabricating or repairing a gasturbine component having a plurality of cooling holes defined therein isprovided. The system includes a robotic device including a slurrydischarge nozzle, and a computing device coupled in communication withthe robotic device. The computing device is configured to determine aparameter of a first cooling hole defined in the gas turbine component,and generate a tool path for forming a protective cap around the firstcooling hole. The tool path is based at least partially on the parameterof the first cooling hole. The computing device is also configured todirect the robotic device to follow the tool path, and direct therobotic device to discharge successive layers of ceramic slurry towardsthe gas turbine component as the tool path is followed such that theprotective cap is formed around the first cooling hole.

In yet another aspect, a non-transitory computer-readable storage mediahaving computer-executable instructions embodied thereon for use infabricating or repairing a gas turbine component having a plurality ofcooling holes defined therein is provided. When executed by at least oneprocessor, the computer-executable instructions cause the processor todetermine a parameter of a first cooling hole defined in the gas turbinecomponent, and generate a tool path for forming a protective cap aroundthe first cooling hole. The tool path is based at least partially on theparameter of the first cooling hole. The computer-executableinstructions also cause the processor to direct a robotic device tofollow the tool path, and direct the robotic device to dischargesuccessive layers of ceramic slurry towards the gas turbine component asthe tool path is followed such that the protective cap is formed aroundthe first cooling hole.

DRAWINGS

These and other features, aspects, and advantages of the presentdisclosure will become better understood when the following detaileddescription is read with reference to the accompanying drawings in whichlike characters represent like parts throughout the drawings, wherein:

FIG. 1 is a schematic illustration of an exemplary gas turbine engine;

FIG. 2 is a block diagram illustrating an exemplary system for use infabricating or repairing a gas turbine component that may be used in thegas turbine engine shown in FIG. 1;

FIG. 3 is a perspective illustration of the system shown in FIG. 2;

FIG. 4 is a cross-sectional illustration of an exemplary gas turbinecomponent having a protective cap formed thereon, in accordance with afirst processing step;

FIG. 5 is a cross-sectional illustration of the gas turbine componentshown in FIG. 4, in accordance with a second processing step;

FIG. 6 is a cross-sectional illustration of the gas turbine componentshown in FIG. 5, in accordance with a third processing step;

FIG. 7 is a schematic illustration of an exemplary layering sequence foruse in forming a protective cap; and

FIG. 8 is a flow diagram of an exemplary method of fabricating andrepairing a gas turbine component.

Unless otherwise indicated, the drawings provided herein are meant toillustrate features of embodiments of the disclosure. These features arebelieved to be applicable in a wide variety of systems comprising one ormore embodiments of the disclosure. As such, the drawings are not meantto include all conventional features known by those of ordinary skill inthe art to be required for the practice of the embodiments disclosedherein.

DETAILED DESCRIPTION

In the following specification and the claims, reference will be made toa number of terms, which shall be defined to have the followingmeanings.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

“Optional” or “optionally” means that the subsequently described eventor circumstance may or may not occur, and that the description includesinstances where the event occurs and instances where it does not.

Approximating language, as used herein throughout the specification andclaims, may be applied to modify any quantitative representation thatcould permissibly vary without resulting in a change in the basicfunction to which it is related. Accordingly, a value modified by a termor terms, such as “about”, “approximately”, and “substantially”, are notto be limited to the precise value specified. In at least someinstances, the approximating language may correspond to the precision ofan instrument for measuring the value. Here and throughout thespecification and claims, range limitations may be combined and/orinterchanged. Such ranges are identified and include all the sub-rangescontained therein unless context or language indicates otherwise.

As used herein, the terms “axial” and “axially” refer to directions andorientations that extend substantially parallel to a centerline of theturbine engine. Moreover, the terms “radial” and “radially” refer todirections and orientations that extend substantially perpendicular tothe centerline of the turbine engine. In addition, as used herein, theterms “circumferential” and “circumferentially” refer to directions andorientations that extend arcuately about the centerline of the turbineengine.

As used herein, the terms “processor” and “computer,” and related terms,e.g., “processing device,” “computing device,” and “controller” are notlimited to just those integrated circuits referred to in the art as acomputer, but broadly refers to a microcontroller, a microcomputer, aprogrammable logic controller (PLC), and application specific integratedcircuit, and other programmable circuits, and these terms are usedinterchangeably herein. In the embodiments described herein, memory mayinclude, but it not limited to, a computer-readable medium, such as arandom access memory (RAM), a computer-readable non-volatile medium,such as a flash memory. Alternatively, a floppy disk, a compactdisc-read only memory (CD-ROM), a magneto-optical disk (MOD), and/or adigital versatile disc (DVD) may also be used. Also, in the embodimentsdescribed herein, additional input channels may be, but are not limitedto, computer peripherals associated with an operator interface such as amouse and a keyboard. Alternatively, other computer peripherals may alsobe used that may include, for example, but not be limited to, a scanner.Furthermore, in the exemplary embodiment, additional output channels mayinclude, but not be limited to, an operator interface monitor.

Further, as used herein, the terms “software” and “firmware” areinterchangeable, and include any computer program storage in memory forexecution by personal computers, workstations, clients, and servers.

As used herein, the term “non-transitory computer-readable media” isintended to be representative of any tangible computer-based deviceimplemented in any method of technology for short-term and long-termstorage of information, such as, computer-readable instructions, datastructures, program modules and sub-modules, or other data in anydevice. Therefore, the methods described herein may be encoded asexecutable instructions embodied in a tangible, non-transitory,computer-readable medium, including, without limitation, a storagedevice and/or a memory device. Such instructions, when executed by aprocessor, cause the processor to perform at least a portion of themethods described herein. Moreover, as used herein, the term“non-transitory computer-readable media” includes all tangible,computer-readable media, including, without limitation, non-transitorycomputer storage devices, including without limitation, volatile andnon-volatile media, and removable and non-removable media such asfirmware, physical and virtual storage, CD-ROMS, DVDs, and any otherdigital source such as a network or the Internet, as well as yet to bedeveloped digital means, with the sole exception being transitory,propagating signal.

Embodiments of the present disclosure relate to turbine engines andrelated systems and methods of fabricating or repairing turbine engines.More specifically, the systems and methods described herein facilitateforming temporary protective caps around cooling holes defined in gasturbine components. The protective caps restrict coating material fromflowing into the cooling holes when a thermal barrier coating is appliedover the gas turbine component. Moreover, the protective caps are formedfrom a ceramic slurry material such that protective caps of varyingshapes and sizes may be formed on the gas turbine component based on thedimensions of each cooling hole. Once the protective caps have cured,the thermal barrier coating is applied over the protective caps, allowedto cure, and the cooling holes are reopened by removing the materialextending over the cooling holes. As such, the thermal barrier coatingis restricted from flowing into the cooling holes, which facilitatesreducing the time and complexity of fabricating or refurbishingcomponents having a thermal barrier coating applied thereto.

FIG. 1 is a schematic illustration of an exemplary gas turbine engine10. In the exemplary embodiment, gas turbine engine 10 includes a gasturbine engine 12 that includes a low pressure compressor 14, a highpressure compressor 16, and a combustor assembly 18 positioneddownstream from high pressure compressor 16. Gas turbine engine 12 alsoincludes a high pressure turbine 20 positioned downstream from combustorassembly 18, a low pressure turbine 22 positioned downstream from highpressure turbine 20, and a power turbine 24 positioned downstream fromlow pressure turbine 22.

In operation, a flow of intake air 26 is channeled through low pressurecompressor 14 and a flow of compressed air is channeled from lowpressure compressor 14 to high pressure compressor 16. The compressedair is discharged from high pressure compressor 16 and channeled towardscombustor assembly 18, where the air is mixed with fuel and combusted toform a flow of combusted gas discharged towards high pressure turbine20. The flow of combusted gas discharged from combustor assembly 18drives high pressure turbine 20 about a centerline 28 of gas turbineengine 12, and the flow of combusted gas is channeled through turbines20 and 22 and then discharged from gas turbine engine 12 in the form ofa flow of exhaust gas 30.

FIG. 2 is a block diagram illustrating an exemplary system 100 for usein fabricating or repairing a gas turbine component 102 that may be usedin gas turbine engine 10 (shown in FIG. 1), and FIG. 3 is a perspectiveillustration of system 100. In the exemplary embodiment, system 100includes a robotic device 104 and a computing device 106 coupled incommunication with robotic device 104. Robotic device 104 includes aslurry discharge nozzle 108. As will be explained in more detail below,slurry discharge nozzle 108 discharges successive layers of ceramicslurry towards gas turbine component 102 for forming a protective cap110 thereon.

Computing device 106 includes a memory 112 and a processor 114,comprising hardware and software, coupled to memory 112 for executingprogrammed instructions. Processor 114 may include one or moreprocessing units (e.g., in a multi-core configuration) and/or include acryptographic accelerator (not shown). Computing device 106 isprogrammable to perform one or more operations described herein byprogramming memory 112 and/or processor 114. For example, processor 114may be programmed by encoding an operation as executable instructionsand providing the executable instructions in memory 112.

Processor 114 may include, but is not limited to, a general purposecentral processing unit (CPU), a microcontroller, a reduced instructionset computer (RISC) processor, an open media application platform(OMAP), an application specific integrated circuit (ASIC), aprogrammable logic circuit (PLC), and/or any other circuit or processorcapable of executing the functions described herein. The methodsdescribed herein may be encoded as executable instructions embodied in acomputer-readable medium including, without limitation, a storage deviceand/or a memory device. Such instructions, when executed by processor114, cause processor 114 to perform at least a portion of the functionsdescribed herein. The above examples are exemplary only, and thus arenot intended to limit in any way the definition and/or meaning of theterm processor.

Memory 112 is one or more devices that enable information such asexecutable instructions and/or other data to be stored and retrieved.Memory 112 may include one or more computer-readable media, such as,without limitation, dynamic random access memory (DRAM), synchronousdynamic random access memory (SDRAM), static random access memory(SRAM), a solid state disk, and/or a hard disk. Memory 112 may beconfigured to store, without limitation, executable instructions,operating systems, applications, resources, installation scripts and/orany other type of data suitable for use with the methods and systemsdescribed herein.

Instructions for operating systems and applications are located in afunctional form on non-transitory memory 112 for execution by processor114 to perform one or more of the processes described herein. Theseinstructions in the different implementations may be embodied ondifferent physical or tangible computer-readable media, such as memory112 or another memory, such as a computer-readable media (not shown),which may include, without limitation, a flash drive and/or thumb drive.Further, instructions may be located in a functional form onnon-transitory computer-readable media, which may include, withoutlimitation, smart-media (SM) memory, compact flash (CF) memory, securedigital (SD) memory, memory stick (MS) memory, multimedia card (MMC)memory, embedded-multimedia card (e-MMC), and micro-drive memory. Thecomputer-readable media may be selectively insertable and/or removablefrom computing device 106 to permit access and/or execution by processor114. In an alternative implementation, the computer-readable media isnot removable.

System 100 also includes a non-destructive inspection device 116 coupledin communication with computing device 106. Non-destructive inspectiondevice 116 is any non-destructive inspection device that enables system100 to function as described herein. Exemplary non-destructiveinspection devices include, but are not limited to, an ultrasonictesting device, an X-ray testing device, and a computed tomography (CT)scanning device. As will be described in more detail below,non-destructive inspection device 116 operates, either continuously orat predetermined intervals, to determine a parameter of cooling holes118 defined in gas turbine component 102. More specifically, at leastone of a size of cooling holes 118, an edge profile of cooling holes118, or a location of cooling holes 118 on gas turbine component 102 aredetermined to facilitate determining the size, shape, and location ofprotective caps 110 to be formed over each cooling hole 118. Forexample, cooling holes 118 of different shapes and sizes may be definedin gas turbine component 102, conducting a non-destructive inspection ofgas turbine component 102 ensures the parameter of each cooling hole 118is accurately determined. In an alternative embodiment, the parameter ofcooling holes 118 is determined from a schematic of gas turbinecomponent 102.

In operation, computing device 106 directs non-destructive inspectiondevice 116 to inspect gas turbine component 102 to determine theparameter of a first cooling hole 120 to be covered. Computing device106 then generates a tool path for forming protective cap 110 aroundfirst cooling hole 120, and the tool path is based at least partially onthe parameter of first cooling hole 120. Computing device 106 thendirects robotic device 104 to follow the tool path, and directs roboticdevice 104 to discharge successive layers of ceramic slurry towards gasturbine component 102 as the tool path is followed such that protectivecap 110 is formed around first cooling hole 118. After first coolinghole 118 is covered, computing device 106 then direct robotic device 104to form protective caps 110 around additional cooling holes based on theparameters determined from the non-destructive inspection describedabove.

In an alternative embodiment, the parameters of first cooling hole 120are determined from a virtual model of gas turbine component 102, orfrom a combination of a non-destructive inspection of gas turbinecomponent 102 and the virtual model of gas turbine component 102.

FIG. 4 is a cross-sectional illustration of an exemplary gas turbinecomponent 102 having protective cap 110 formed thereon, in accordancewith a first processing step, FIG. 5 is a cross-sectional illustrationof gas turbine component 102, in accordance with a second processingstep, and FIG. 6 is a cross-sectional illustration of gas turbinecomponent 102, in accordance with a third processing step. In theexemplary embodiment, and as described above, computing device 106(shown in FIG. 3) generates a tool path for forming protective cap 110around first cooling hole 120 after at least one parameter of firstcooling hole 120 is determined. More specifically, computing device 106generates a hollow three-dimensional representation that definesprotective cap 110, and determines an arrangement of a plurality ofindividual layers 122 for forming the three-dimensional representationof protective cap 110. The plurality of individual layers 122 correspondto the successive layers of ceramic slurry used to form protective cap110 on gas turbine component 102.

The three-dimensional representation of protective cap 110 is generatedin accordance with one or more design parameters. Exemplary designparameters include, forming protective cap 110 from a base portion 124and a dome portion 126, forming base portion 124 to follow the contourof side edges of first cooling hole 120, forming base portion 124 havinga thickness substantially equal to a thickness of a thermal barriercoating layer 128 (shown in FIG. 5) to be applied over protective cap110, forming dome portion 126 with a converging shape to facilitatecovering first cooling hole 120, and forming dome portion 126 such thata vertical profile thereof follows a circular arc.

Computing device 106 generates the tool path for forming protective cap110 by slicing the three-dimensional representation of protective cap110 into the plurality of individual layers 122. The thickness of eachindividual layer 122, and the percentage overlap between adjacentindividual layers 122 may be modified along the height of protective cap110 such that protective cap 110 is formed in a time-efficient mannerwith enhanced stability. More specifically, in one embodiment, computingdevice 106 defines a progressive reduction in thickness of the pluralityof individual layers 122 as a distance between each individual layer 122and gas turbine component 102 increases. In another embodiment,computing device 106 defines a progressive reduction in overlap betweenadjacent individual layers 122 as a distance between the adjacentindividual layers 122 and gas turbine component increases.Alternatively, individual layers 122 may have a uniform thickness, andthe overlap between adjacent individual layers 122 is uniform along theheight of protective cap 110.

In the exemplary embodiment, base portion 124 is formed from at least afirst individual layer 130 applied directly onto gas turbine component102, and dome portion 126 is formed from successive second individuallayers 132 applied over first individual layer 130. The layers ofprotective cap 110 may be formed from any material that enables system100 to function as described herein. An exemplary material includes, butis not limited to, an aqueous ceramic-based slurry formed with aninorganic binder material. As such, the slurry has self-curingcharacteristics, and the inorganic binder material is generallyresistant to burnout during one or more curing cycles. The inorganicbinder material may also be a refractory cement capable of withstandingincreased temperatures during deposition of thermal barrier coatinglayer 128 thereon, and also of the high temperatures during operation ofgas turbine engine 10 (shown in FIG. 1). The slurry may also include arefractory ceramic material, such as yttria-stabilized zirconia (YSZ)powder with a multimodal size distribution. The bimodal sizedistribution enables larger solids loading, and sub-micron sized powderis used for rheology control.

In one embodiment, first individual layer 130 is fabricated from a firstmaterial, and the successive second individual layers 132 are fabricatedfrom a second material different from the first material. The firstmaterial is a bond coat layer, and the second material is the ceramicslurry described above. The bond coat layer facilitates defining asmooth transition between gas turbine component 102 and secondindividual layers 132, and also provides enhanced adhesion and corrosionprotection to gas turbine component 102. The bond coat layer may beformed from any suitable material such as, but not limited to, acomposition including chromium, aluminum, yttria, and at least one ofnickel, cobalt, or a combination thereof.

Referring to FIG. 5, a bond coat layer 134 and a thermal barrier coatinglayer 128 are applied over gas turbine component 102 and over protectivecap 110 after the ceramic slurry has cured. Referring to FIG. 6, firstcooling hole 120 is cleared by removing thermal barrier coating layer128, bond coat layer 134, and at least a portion of protective cap 110.The material is removed by grinding or polishing, for example. As such,protective cap 110 restricts bond coat layer 134 and thermal barriercoating layer 128 from entering first cooling hole 120 during theapplication process. Bond coat layer 134 may be formed from any suitablematerial such as, but not limited to, a composition including chromium,aluminum, yttria, and at least one of nickel, cobalt, or a combinationthereof.

FIG. 7 is a schematic illustration of an exemplary layering sequence foruse in forming protective cap 110. In the exemplary embodiment, and asdescribed above, the plurality of individual layers 122 progressivelyreduce in thickness as a distance from gas turbine component 102 (shownin FIG. 3) increases, and the amount of overlap between adjacentindividual layers 122 progressively reduces as the adjacent individuallayers 122 are positioned a greater distance from gas turbine component102. More specifically, the amount of overlap to define between adjacentindividual layers 122 is determined based on the thickness of eachsuccessive individual layer 122. As such, thicker individual layers 122are layered closer to gas turbine component 102 to reduce the number oflayers required to form protective cap 110, and thinner individuallayers 122 are layered farther away from gas turbine component 102 thanthe thicker individual layers 122 to ensure protective cap 110 remainsstable as the ceramic slurry cures.

An exemplary technical effect of the system and methods described hereinincludes at least one of: (a) restricting coating material from enteringcooling holes defined in a component; (b) reducing the time andcomplexity required to clear cooling holes of coating material; and (c)enabling the formation of protective caps of varying shapes and sizes inan automated manner.

Exemplary embodiments of a turbine engines and related components aredescribed above in detail. The system is not limited to the specificembodiments described herein, but rather, components of systems and/orsteps of the methods may be utilized independently and separately fromother components and/or steps described herein. For example, theconfiguration of components described herein may also be used incombination with other processes, and is not limited to practice withonly turbine engines and related methods as described herein. Rather,the exemplary embodiment can be implemented and utilized in connectionwith many applications where adaptably forming three-dimensionalstructures is desired.

Although specific features of various embodiments of the presentdisclosure may be shown in some drawings and not in others, this is forconvenience only. In accordance with the principles of embodiments ofthe present disclosure, any feature of a drawing may be referencedand/or claimed in combination with any feature of any other drawing.

Some embodiments involve the use of one or more electronic or computingdevices. Such devices typically include a processor or controller, suchas a general purpose central processing unit (CPU), a graphicsprocessing unit (GPU), a microcontroller, a reduced instruction setcomputer (RISC) processor, an application specific integrated circuit(ASIC), a programmable logic circuit (PLC), and/or any other circuit orprocessor capable of executing the functions described herein. Themethods described herein may be encoded as executable instructionsembodied in a computer readable medium, including, without limitation, astorage device and/or a memory device. Such instructions, when executedby a processor, cause the processor to perform at least a portion of themethods described herein. The above examples are exemplary only, andthus are not intended to limit any way the definition and/or meaning ofthe term processor.

This written description uses examples to disclose the embodiments ofthe present disclosure, including the best mode, and also to enable anyperson skilled in the art to practice embodiments of the presentdisclosure, including making and using any devices or systems andperforming any incorporated methods. The patentable scope of theembodiments described herein is defined by the claims, and may includeother examples that occur to those skilled in the art. Such otherexamples are intended to be within the scope of the claims if they havestructural elements that do not differ from the literal language of theclaims, or if they include equivalent structural elements withinsubstantial differences from the literal languages of the claims.

What is claimed is:
 1. A method of fabricating or repairing a gasturbine component having a plurality of cooling holes defined therein,said method comprising: determining a parameter of a first cooling holedefined in the gas turbine component; generating a tool path for forminga protective cap around the first cooling hole, the tool path based atleast partially on the parameter of the first cooling hole; directing arobotic device to follow the tool path; and discharging successivelayers of ceramic slurry towards the gas turbine component as the toolpath is followed such that the protective cap is formed around the firstcooling hole.
 2. The method in accordance with claim 1, whereindetermining a parameter comprises determining at least one of a size ofthe first cooling hole, an edge profile of the first cooling hole, or alocation of the first cooling hole on the gas turbine component.
 3. Themethod in accordance with claim 1, wherein determining a parametercomprises conducting a non-destructive inspection of the gas turbinecomponent.
 4. The method in accordance with claim 1, wherein generatinga tool path comprises determining an arrangement of a plurality ofindividual layers for forming a three-dimensional representation of theprotective cap, wherein the plurality of individual layers substantiallycorrespond to the successive layers of ceramic slurry.
 5. The method inaccordance with claim 4, wherein determining an arrangement of aplurality of individual layers comprises defining a progressivereduction in thickness of each individual layer of the plurality ofindividual layers as a distance between each individual layer and thegas turbine component increases.
 6. The method in accordance with claim5, wherein discharging successive layers of ceramic slurry comprisescontrolling at least one of a flow rate for discharging the ceramicslurry or a tool speed to define a progressive reduction in thickness ofthe successive layers of ceramic slurry that corresponds to theprogressive reduction in thickness of each individual layer of theplurality of individual layers.
 7. The method in accordance with claim4, wherein determining an arrangement of a plurality of individuallayers comprises defining a progressive reduction in overlap betweenadjacent individual layers as a distance between the adjacent individuallayers and the gas turbine component increases.
 8. The method inaccordance with claim 1, wherein discharging successive layers ofceramic slurry comprises: applying a first layer of ceramic slurry onthe gas turbine component, the first layer fabricated from a firstmaterial; and applying successive second layers of ceramic slurry on thefirst layer, the successive second layers fabricated from a secondmaterial different from the first material.
 9. The method in accordancewith claim 8, wherein the first material is formed from a compositionincluding chromium, aluminum, yttria, and at least one of nickel,cobalt, or a combination thereof.
 10. The method in accordance withclaim 8, wherein the second material is an aqueous ceramic-based slurrywith an inorganic binder material.
 11. The method in accordance withclaim 10, wherein the inorganic binder material comprises a refractorycement.
 12. The method in accordance with claim 1, wherein the ceramicslurry includes yttria-stabilized zirconia powder with a multimodal sizedistribution.
 13. The method in accordance with claim 8, furthercomprising: applying a dome portion over the successive second layers,wherein the first layer, the successive second layers, and the domeportion have a profile that follows a circular arc.
 14. The method inaccordance with claim 10, further comprising: curing the ceramic slurry.15. The method in accordance with claim 14, further comprising: applyinga bond coat layer over the protective cap after curing the ceramicslurry.
 16. The method in accordance with claim 15, wherein the bondcoat layer is formed from a composition including chromium, aluminum,yttria, and at least one of nickel, cobalt, or a combination thereof.17. The method in accordance with claim 15, further comprising: applyinga thermal barrier coating layer over the bond coat layer.
 18. The methodin accordance with claim 17, further comprising: clearing the firstcooling hole by removing the thermal barrier coating layer, the bondcoat layer, and at least a portion of the protective cap.
 19. The methodin accordance with claim 18, wherein removing comprises grinding orpolishing.